During reentry into the atmosphere, a vehicle is subjected to extreme thermal conditions. As the vehicle contacts the atmosphere at very high speeds, frictional forces release high levels of thermal energy which can raise the temperature to levels which are destructive to the outer shell. To protect the vehicle from high temperatures and wind shear, the vehicle's outer shell is typically covered with TPMs, which act as insulators and are designed to withstand these extreme thermal conditions.
Carbon-carbon (C—C) composites are one class of TPM which have been employed under such conditions with proven effectiveness. The success of a particular TPM requires that the system have sufficient mechanical strength at high temperatures, produce endothermic reactions upon decomposition, and have a high surface emissivity.
In its simplest form, a carbon-carbon composite is manufactured by combining carbon fibers with an organic resin, usually a high carbon yield epoxy or phenolic resin, and the resulting carbon fiber and resin matrix cured to achieve a three dimensional structure such as a tile, billet or other object. The matrix has a density, a void volume and a degree of mechanical strength.
The carbon fiber and resin matrix is then subjected to a high temperature treatment which decomposes the resin matrix to pure carbon, a process called charring or carbonization. Charring changes the resin coating from an organic resin to free carbon which coats the carbon fibers and partially fills the void spaces of the matrix with free carbon. The TPM may be subjected to several charring cycles, a process known as densification. The result of densification is to create a more rigid substrate, with a decreased void volume. The char surface of the substrate has a high temperature structural capability, which is a desirable characteristic.
Conventional C—C composites are manufactured in such a way so as to produce a highly filled and rigidized structure with a minimum of porosity. There are many ways for C—C materials to be densified including infiltration with petroleum pitch, impregnation with phenolic or other organic resins, or carbon vapor infiltration (CVI) using low molecular weight hydrocarbons such as methane. Any substance used for densification should have a high carbon char yield. Repeated cycles of impregnation and carbonization are required to first infuse the material with the carbon materials and then to heat them to a sufficiently high temperature (generally above 500° C.) to char the infiltrant and to create porosity for further densification cycles. A typical density range for a C—C composite with 5% porosity is approximately 1.6 to 1.8 g/cc, depending on the infiltrants and carbon fibers used in the composite.
The use of C—C composite TPMs on long duration, high altitude hypersonic reentry vehicles exhibit, however, some characteristics which can severely restrict mission performance. A major limitation of these materials is that they are subject to oxidation at extreme thermal conditions. The oxidation that these TPMs experience during long duration reentry can result in large shape changes to the vehicle aero-shell. Shape changes that adversely affect the mechanical strength and aerodynamics of the vehicle are unacceptable. To compensate for the loss of mechanical or structural integrity, which can lead to shape changes, typically the thickness of the material is increased. Increasing the thickness, however, adds unacceptably to the weight and volume of the vehicle, thus reducing the payload capacity and increasing cost.
While the C—C class of TPMs make them good candidates for aerospace applications due to their excellent high temperature structural properties oxidation shape changes can still be a problem. To address this, extensive efforts have been expended on oxidation resistant coatings for C—C composites with, however, limited success. The coatings developed to date are restricted to temperature levels generally below those experienced during reentry into the atmosphere, or in other high temperature applications. Also, coating costs and durability (durability in the form of handling microcracking, the occurrence of pinholes, particle impacts and damage from ground handling) are serious issues when one is considering coatings for use on C—C composite TPMs.
Ablation technology employs several mechanisms to manage the high levels of thermal energy released during reentry. Three of these are the vaporization and decomposition (pyrolysis) of the resin and subsequent transpirational cooling at the boundary layer. All of these processes absorb heat. Producing large amounts of gas is one measure of an ablation based system's ability to absorb heat. The production of gas can also be increased by impregnating the C—C substrate with an organic material specifically designed to vaporize and pyrolyze upon exposure of the system to high heat loads. Materials used in these passive transpiration systems, known as coolants, include materials such as polyethylene or epoxy, acrylic or phenolic resins.
Under such a system, there is created within the material a pyrolysis zone, where the resin and any supplemental coolants present are heated to temperatures where the organic materials decompose. The effect is the absorption of heat and the creation of additional carbon which can remain in the pyrolysis zone and/or be deposited on the carbon fibers and within the void volume of the substrate. Thus, the C—C ablator's final weight and ability to absorb heat are directly related to the amount of available resin in the C—C composite prior to reentry.
At the surface of the C—C ablator, heat is reradiated due to the refractive properties of the carbon substrate. In addition, the gasses produced in the pyrolysis zone within the C—C ablator are released to the surface at a relatively cool temperature when compared to the conditions at the surface. This effect, known as pyrolysis gas transpiration, provides cooling at the surface of the TPM. The disadvantages of the passive transpiration systems described herein include the high overall density of the material and the high internal pressure cause by the sudden buildup of gasses within the material. Ablation systems which can create and then release large volumes of gas thus exhibit a greater capacity to absorb and dissipate the heat of reentry.
In this regard, the structure of the C—C substrate is important to the overall effectiveness of the ablator. The void volume can be filled with a resin or other coolant to provide the raw material for production of gasses. In addition, methods of construction of the substrate can allow for greater transpiration pathways for release of the gasses. Systems which generate large volumes of gas over a short period of time also generate high internal pressures. Such pressure causes internal cracking in the substrate (microcracks) and also spalling at the surface. These effects are destructive to the mechanical integrity of the system and can lead to system failure. Therefore, improved transpiration pathways also protect the system from the effects of this internal pressure.
U.S. Pat. No. 5,635,300 to Kostikov, et. al., describes an advancement in the art of C—C or ceramic ablators through the introduction of silicon based resins to the C—C substrate. Upon decomposition and subsequent exposure to the very high temperatures at the surface, the silicon resin reacts with the carbon substrate to form a silicon carbide (SiC) coating on those fibers experiencing the high temperature conditions. The formation of SiC is more resistasnt to oxidation than carbon, and thus acts to strengthen the carbon substrate by forming a SiC skeleton in the areas of extreme temperature. When prolonged conditions of high temperature and wind shear at the surface lead to loss of SiC, the newly exposed carbon substrate undergoes further reaction to form new SiC, thus regenerating the protective skeleton.
The SiC layer which forms over the fibers of the carbon substrate in the interior of the ablator has a different coefficient of thermal expansion (CTE) than the carbon itself. The result is that when the system is subjected to temperature changes, the SiC coating within the carbon substrate forms microcracks. These cracks form passageways for the entry of air, which leads to oxidation of the carbon substrate, with resultant loss of strength and integrity of the ablator.
In Kostikov, a carbon-SiC substrate is created by obtaining a carbon-plastic preform composed of carbon fibers and a thermosetting resin binder and heat treating to form a coke matrix reinforced by carbon fibers. The coke matrix is densified by infiltrating with pyrolytic carbon and heat treating the preform at 1900 to 2000 deg. C. According to this invention, pore channels are formed following crystalization of the carbon deposited upon the matrix. Densification is followed by treatment with silicon which forms a SiC skeleton in the pore spaces of the composite. The carbon fibers can be in the form of a woven fabric or woven substrate.
U.S. Pat. No. 5,672,389 to Tran, et. al., discloses a low density ceramic ablator which employs a fibrous ceramic substrate which has, prior to impregnation with a resin matrix, a density of about 0.15 to 0.2 g/cc. Tran includes carbon fibers within a definition of the term ceramic. The ceramic substrate is impregnated with a low viscosity solution containing of an organic resin in a solvent. The excess infiltrant is removed, followed by removal of the solvent under vacuum, leaving resin-coated fibers, and a substrate having an average density of from 0.15 to 0.4 g/cc. Tran discloses that the resulting ablator may have either a uniform distribution of resin on the ceramic fibers, or a non-uniform distribution. The non-uniform distribution has the benefit of achieving the necessary degree of ablation at the outer surface, while being lightweight at the inner surface, where extreme temperatures are not experienced.
In addition, at the oxidizing conditions at the surface, silicon reacts with atmospheric oxygen to form a coating of silicon oxide (SiO2), which manifests itself as a glassy layer on the outer surface of the ablator. This SiO2 plus free carbon and free SiC mixture has a high surface emissivity which improves the ability of the material to radiate heat from the surface due to convection and re-radiation from the carbon substrate.
U.S. Pat. No. 5,965,266 to Goujard, et. al. discloses a carbon-SiC (C—SiC) composite TPM which has a self-healing mechanism for in-situ repair of the C—SiC matrix. The matrix is heat treated to form SiC and boron carbide (BC) over the C—SiC matrix. The SiC layer improves the mechanical strength of the system. However, due to the difference in CTE, the matrix experiences destructive cracking when exposed to the temperature changes experienced during reentry. These cracks create pathways that allow the entry of air, which causes oxidation of the C—SiC matrix, thus mechanically weakening the structure of the TPM.
The invention of Goujard provides available free silicon and boron as glass precursors which react with the exposed carbon under the high temperatures and oxidizing conditions of reentry. The glass precursors to form a self healing glass layer within the cracks, closing the pathway for internal oxidation of the substrate.
In addition to the strength and density of an ablation system, the manner in which the material is mounted onto the aero-shell, and the incorporation of additional layers of insulation also have bearing of the success of the TPM. U.S. Pat. No. 3,152,548 to Schwartz discloses a system whereby a series of coiled wires are attached to the aero-shell and the ceramic TPM is mounted onto the metal coils, thus creating a space between the aero-shell and the ceramic TPM. This space is filled with a pliable thermal insulating material, thus providing added insulation protection to the aero-shell. The patent discloses that the use of coiled wire compensates for differences in thermal expansion between the metal aero-shell and the ceramic insulator.